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Topic: Basic Rocket Science Q & A (Read 298110 times)

Thanks for the thoughtful remarks Isaac. The way you express the math is very clear.

You're welcome!

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As for elliptical orbits versus LaGrange points: I don't see why you think the "wait time" is going to be a big deal--after all, EML1 or 2 is only a week or so from Earth. Since Mars gravity is so weak compared to the Sun, the Sun-Mars L1 point is going to be fairly close to Mars. Also, Lagrange points are ideal locations for propellant depots. (Not to mention that EML1/2 is close to Lunar propellant sources ).

If you want to minimize delta-v, the wait period will be on the same order of magnitude as the orbital period. You use multiple orbits to reduce the delta-v and take advantage of the Lagrange point instability.

For EML1 and EML2 the wait isn't such a big deal because the orbital period is only a month. For Sun-Mars L1 and L2, the orbital period is almost 2 years. Yes, you can get to/from Sun-Mars L1 and L2 faster by using more delta-v--but this defeats the purpose. You might as well have staged at any random location outside Mars's Hill sphere.

Also, the direction of the hyperbolic escape trajectory after leaving Mars won't line up well with the desired direction to get back to Earth. This isn't a problem for EML1/EML2 because of the way the Moon orbits Earth. Within a month, it will be in the correct alignment for the desired Earth escape trajectory. But the Sun-Mars lagrange point retain the same orientation with respect to Mars's orbit at all times.

As for the ideal location for propellant depots--I favor placing tankers in highly elliptical orbits rather than lagrange points. It maximizes the delta-v advantage while reducing the wait times.

For a given Isp (I don't care if it's 50 or 5000) and in-space trajectory change (orbit change, maneuvers, escapes), does high thrust or low thrust end up with less propellant required?

Here's a quick math-like "proof" that high thrust always uses at most the delta vee of low thrust: a high thrust engine can simulate a low thrust engine by operating in a series of short pulses.

You can sometimes also go the other way simulate a high thrust engine with a low thrust engine, at least in theory. For example you can presumably simulate a high-thrust Hoffmann transfer using a low thrust engine by firing the engine for a short pulse at perigee over many orbits until apogee is raised, and then short pulses at apogee. This trick is of course limited to circumstances where you have patience and repeated orbits.

More practically high thrust enables better use of the Oberth effect, which can reduce delta vee. For example a low-thrust spiral from low earth orbit to escape uses 100% of orbital speed worth of delta vee whereas an infinite-thrust impulse uses only 41%. (See http://en.wikipedia.org/wiki/Hohmann_transfer_orbit .)

Of course propellant usage is more complicated than just delta vee, but surely you can't expect someone answering your short question to incorporate tradeoffs between dry mass and delta vee!

For a given Isp (I don't care if it's 50 or 5000) and in-space trajectory change (orbit change, maneuvers, escapes), does high thrust or low thrust end up with less propellant required?

It really depends... Given that rocket engines are limited in their T/W, a "really high"-thrust stage will take more propellant than a "moderately high"-thrust, since the dry mass that needs to be put through your delta-v will be higher for the high-thrust stage. Also, higher thrust (if higher than launch accelerations, etc) can also mean higher structural mass requirements in both the rocket stage and payload, thus increasing total propellant needed.

And, of course, if you have too low thrust, your long burn means you can't burn all the necessary propellant deep in the gravity well, thus you need more propellant for a given escape velocity, etc.

But with multiple passes and multiple burns (raising apogee, not perigee, say for escape), it can be possible to "simulate" a higher thrust rocket engine with a much lower thrust rocket engine...

Chris Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

For a given Isp (I don't care if it's 50 or 5000) and in-space trajectory change (orbit change, maneuvers, escapes), does high thrust or low thrust end up with less propellant required?

It really depends... Given that rocket engines are limited in their T/W, a "really high"-thrust stage will take more propellant than a "moderately high"-thrust, since the dry mass that needs to be put through your delta-v will be higher for the high-thrust stage. Also, higher thrust (if higher than launch accelerations, etc) can also mean higher structural mass requirements in both the rocket stage and payload, thus increasing total propellant needed.

And, of course, if you have too low thrust, your long burn means you can't burn all the necessary propellant deep in the gravity well, thus you need more propellant for a given escape velocity, etc.

But with multiple passes and multiple burns (raising apogee, not perigee, say for escape), it can be possible to "simulate" a higher thrust rocket engine with a much lower thrust rocket engine...

I believe this actually was done to save a com sat using the RCS because the apogee engine failed but it's also an advantage VASIMR has over ion engines.A VASIMR rocket is not bothered as much by multiple restarts as a gridded ion engine is.

Might add a another law to the list if a change only nets an extra 3% performance gain but increases the cost by 15% it's best not to implement it.I'll call it Korolev's law as Russians LVs often tend just accept a lower payload mass fraction if a given design element is lower cost or more robust.

Russian LVs tend to be simple and robust while western LVs tend to be highly tuned hotrods.

...A VASIMR rocket is not bothered as much by multiple restarts as a gridded ion engine is....

...And the main reason is because VASIMR has never flown.

I actually kind of doubt this. Have you seen the flight profile of Dawn (which uses a gridded ion engine)? It's all chopped up... Very many starts and stops.

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Chris Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

I believe this actually was done to save a com sat using the RCS because the apogee engine failed but it's also an advantage VASIMR has over ion engines.A VASIMR rocket is not bothered as much by multiple restarts as a gridded ion engine is.

Might add a another law to the list if a change only nets an extra 3% performance gain but increases the cost by 15% it's best not to implement it.I'll call it Korolev's law as Russians LVs often tend just accept a lower payload mass fraction if a given design element is lower cost or more robust.

Russian LVs tend to be simple and robust while western LVs tend to be highly tuned hotrods.

Yeah, it was on Advanced EHF, a military comsat. Apogee failed, so they had to use the RCS to put it in a stable orbit, otherwise it would have deorbited very quickly. It's been spiraling for a year on Hall thrusters since, which also don't care about restarts, and actually are flying .

One rookie question: Do rocket stages using hypergolic fuel have the danger of fuel/oxidizer sloshing? Why?

Yes they do. If you launched a tank with water, it would too. Slosh is nothing exotic, it can just lead to hard to predict effects depending on how it couples with the vehicle. If you have a liquid, it can slosh.

Let's try to assume a realistic rocket that uses methane as a propellant. No oxygen for this rocket.

How much methane is needed to raise 100 kilograms of cargo onto LEO (400 kilometers above the surface)?

How much power and energy is needed to heat up that methane and produce enough thrust to raise that mass of 100 kilograms onto LEO (400 km)?

Well, this is a very strange rocket, and I'm not totally sure of what you're asking, but I gave it a shot. I've attached a file. Basically, we're limited by materials to a temp around combustion temp anyway. So, I've limited to 3000K. Hot methane has a specific heat ratio of about 1.2, molecular weight of 16. Assumed a nozzle with a pressure ratio of 100, which is reasonable (and going to 10 or 1000 doesn't change it that much). So, the Isp is only 314s under these assumptions.

To get a delta-V of 9600 m/s (Low Earth Orbit w/ g losses, drag), and assuming a weightless structure (any reasonable structure, and it won't get to LEO on an SSTO), you need 2160kg of methane.

The energy required is a function of the total propellant mass, and the exit velocity (Isp), I get 10.2 GJ, or about the energy consumed by a typical American car in 2 months.

The max power will occur at max thrust (if Isp is constant). Thrust should be sized by thrust to weight ratio, I chose 1.1 at liftoff. Power will be 88.4 MW at that time, or about the power of a GE90 jet engine (don't get any ideas, different tech).

Might add a another law to the list if a change only nets an extra 3% performance gain but increases the cost by 15% it's best not to implement it.I'll call it Korolev's law as Russians LVs often tend just accept a lower payload mass fraction if a given design element is lower cost or more robust.

Russian LVs tend to be simple and robust while western LVs tend to be highly tuned hotrods.

My uncle was a F-86 pilot in the Korean War and tells a story about the MIG-15's. They were really rugged, built like tanks and not swiss watches. He talks about one that he shot down. It was the last of his rounds so he was starting to head home (he thought he had a kill) when he noticed that the pilot was actually landing the plane in a mushy field. The pilot got out and did a quick field repair to the rudder, where my uncle had hit him, and got back in the cockpit and took off again from a waterlogged field, actually a rice paddy. Positionally, my uncle had the tactical advantage so he was able to get right behind the MIG pilot. That pilot knew he was dead (but he didn't know my uncle was out of shells). My uncle was so impressed with the aircraft and what that pilot did that he wagged him and then turned for home. He took a hell of a chance that the MIG wouldn't turn to engage but it didn't. The MIG pilot must have respected my uncle as much as my uncle respected him.

Just last year I was at an air show in Cheyenne, Wyoming and watched a MIG-15 perform. Afterward, I got to sit in the MIG and the pilot explained all the instrumentation and controls to me. Pretty sparse, only what the pilot really needed. The MIG was privately owned. The owner-pilot belongs to a flying club of vintage fighter aircraft and tours the country in the summer with his MIG. That MIG actually saw combat in Korea, which makes it 60 years old.

I relay that story to make the point that the Russians have been building things like that for decades. They know where the point of diminishing returns is and they don't cross it – they can't afford to. What they do do is make their machines as rugged (not sophisticated) as possible within the budget available. NASA could learn a few things from them in that regard.

« Last Edit: 08/19/2011 02:24 PM by clongton »

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Chuck - DIRECT co-founderI started my career on the Saturn-V F-1A engine

Another rookie question. If you'd have a tank pressurized over the triple point of a liquid (let's use H2 for the example), what would happen if you start with a full liquid charge?Would it allow better density? (disregarding the tank's mass, of course)And if you went beyond the critical point?

If you'd have a tank pressurized over the triple point of a liquid (let's use H2 for the example), what would happen if you start with a full liquid charge?Would it allow better density? (disregarding the tank's mass, of course)And if you went beyond the critical point?

I assume you're familiar with slush hydrogen and densified LOX, but those are more from lower temperature than higher pressure. The former is advantageous because higher pressure requires heavier tanks.

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If I like something on NSF, it's probably because I know it to be accurate. Every once in a while, it's just something I agree with. Facts generally receive the former.

A modern ICBM is able to hit multiple targets using multiple independently targetable reentry vehicles with an accuracy of 150 meters and very high reliability. Most importantly these ICBMs are fire and forget. They only require a destination. On top of that the launch team for these rockets are only a few people.

My question is why do modern rockets require such large launch teams? The Shuttle has a launch team of 500 people.